Rocket engines can be divided into four main categories consisting of ignition combustion engines, catalytic decomposition engines, pulsed detonation engines, and solid propellant engines. Ignition combustion engines are controlled by opening liquid propellant control valves and allowing propellants to flow into a combustion chamber. The combustion is either self-initiated due to a hypergolic reaction between propellants or by another ignition source, such as a spark plug. After combustion has begun, thrust is produced until the propellant control valves are closed and flow of propellant into the combustion chamber has stopped. The thrust level produced by combustion is directly related to the flow rate of propellants into the combustion chamber and therefore the pressure in the propellant tanks that is driving the propellant flow while the propellant control valves are open. The thrust produced is proportional to the feed pressure in the propellant tank and the flow rate into the combustion chamber. The propellant flow stops when the control valves are closed and thrust is then stopped. Operations with rocket combustion require rapid ignition and high combustion temperature. If ignition delay is too slow, unburned propellant will be exhausted from the rocket engine, reducing performance and potentially causing damage to the vehicle. If combustion temperature is too high, materials within the combustion chamber will be degraded resulting in catastrophic failure of the engine.
A catalytic decomposition engine uses a monopropellant that flows into a decomposition chamber and produces thrust as long as the control valve is open and the monopropellant is flowing into the decomposition chamber. Referring to FIG. 2, a bipropellant ignition combustion engine and the monopropellant catalytic decomposition engine function as rocket propulsion systems where thrust is developed during control valve actuation when the propellant is flowing into the combustion chamber and thrust is terminated when the control valve is closed, in an on and off cyclic operation.
Pulse detonation engines rely a fixed quantity of two propellants, including fuel and an oxidizer, that are fed into the combustion chamber in pulses by concurrently opening two propellant control valves and then concurrently closing the two propellant control valves. When the control valves are closed, the flow of propellant into the chamber is stopped and the chamber contains the desired quantity of propellant. Combustion is had by either self-initiation by a hypergolic reaction or by another ignition source. Upon ignition with the control valve closed, the fuel and the oxidizer propellants combust, resulting in a supersonic detonation wave expanding from the combustion chamber and exiting the engine through a diverging exhaust nozzle. Once the detonation wave has exited the combustion chamber through the exhaust nozzle, the control flow, ignition, and combustion processes are repeated. An integral feature of pulsed detonation engines is a large pressure spike in the combustion chamber when detonation occurs. This pressure spike creates design problems due to structural concerns with the combustion chamber and the transmission of the shock to the rest of the vehicle. Current pulse detonation engines do not use catalytic decomposition due to the requirement for rapid ignition to generate a detonation. If a detonation could be generated by a catalyzed reaction, the force of the shock could damage the catalyst bed. As shown in FIG. 2, the pulse detonation wave occurs just after closing the control valve.
Solid propellant engines use solid propellants to create thrust in a continuous mode without the use of fuel control valves. Throat exhaust valves in converging and diverging exhaust nozzles have been particularly used in the solid propellant engines, often referred to as rocket motors, where the throat valve is used to throttle the rocket motor, that is, to raise and lower the exhaust flow for controlling the thrust on demand, because the solid propellant engine cannot otherwise be controlled. Thrust is controlled because the valve reduces the throat area and restricts the exhaust flow. A solid propellant engine cannot completely close the throat exhaust valve because the combustion will either be completely extinguished and not reignited, or will accelerate resulting in a combustion chamber overpressure and rupture. Liquid propellant engines, including the ignition combustion engines and the catalytic decomposition engines, do not use throat exhaust valves because thrust can controlled more easily through the use of the liquid propellant-flow control valve.
Gas generator propulsion systems use throat exhaust valves. In a monopropellant gas generator propulsion system, liquid propellant flows into a catalyst bed and decomposes into a warm gas and exhausts into a warm gas storage tank. The warm gas can then be stored and exhausted when thrust is needed by flowing the warm gas through a manifold to any number of gas thrusters. The gas thrusters are a valve with a throat and a nozzle located at the exit of the valve. The valve acts as a throat exhaust valve because the valve controls the flow of decomposed propellant through the throat and out the exhaust nozzle. The warm gas storage tank is used to hold the decomposed gas for later time released usage, in hours, days, months, or even years.
Pressurized gas propulsion systems have been used on existing picosatellites. Recently, picosatellite sized vehicles have been proposed for numerous space missions. However, total available ΔV from the propulsion system has limited the orbits available to these vehicles. With insufficient ΔV, the vehicles would be unable to perform sufficient deorbit maneuvers and could eventually create a hazard for other space vehicles. The pressurized gas propulsion system is safe to use and requires only a simple exhaust system, but has low performance capability. Pressurized gas systems can be designed to store the gas propellant as a high-pressure compressed gas or as a saturated liquid. Both storage methods have advantages and disadvantages. The primary trade exists between the relatively high specific impulse (Isp) efficiency provided by compressed gas versus the high storage density obtained by saturated liquids. Two examples are discussed. A 5.0 in3 volume of Nitrogen at 800 psia operating at Isp=65 will result in a total impulse of 0.746 lbf*sec=3.32 Ns. With an equal volume, Butane as a saturated liquid will have a pressure of 31 psia, and with an operating Isp=40, a total impulse of 4.2 lbf*sec=18.6 Ns will result. The large increase in delivered impulse results from the significant density improvement with a saturated liquid despite having a lower specific impulse. Further large gains in delivered impulse could be achieved with a liquid monopropellant by maintaining the high density but simultaneously achieving a high specific impulse. Four cubic inches of hydrazine with 1 in3 gas volume for pressurization and with a specific impulse of 200 lbf*sec/lbm a total impulse of 29 lbf*sec=129 Ns will result. Despite this performance advantage, hydrazine has not been used for several reasons. The primary concern is with hydrazine toxicity that creates severe handling and test restrictions that can add significant cost to the overall system. These same handling concerns create safety issues when integrating the picosatellite with the primary payload on the launch vehicle. In a typical hydrazine satellite thruster, propellant is introduced into the combustion chamber by opening a solenoid valve and allowing liquid flow until thrust is terminated by closing the valve. Ignition delay times are less then 5.0 ms and time to 90% thrust is typically less then 50.0 ms. With this rapid ignition and pressurization of the chamber, combustion occurs mainly as a steady process even with very short pulse durations. This provides accurate control and handling for sensitive satellite systems.
In contrast to hydrazine, hydroxyl ammonium nitrate (HAN) is generally considered to be a non-toxic, that is, a green propellant. HAN usage minimizes the handling concerns raised by hydrazine. In addition to low toxicity, HAN is a stable liquid thereby further improving the handling characteristics. In many respects, HAN could be considered safer than butane and compressed gas due to the lower flammability of HAN in air and the reduced pressure for storage of HAN. The largest drawback of HAN usage is the difficulty rocket engines have had in achieving reliable, repeatable combustion ignition at reasonable preheat temperatures with catalysts that do not melt. HAN ignition delay is too long for current conventional rockets and the HAN decomposition temperature is too high for current conventional catalysts. Monopropellant catalytic decomposition engines for satellite propulsion systems using HAN cannot achieve short ignition delay times because the initial decomposition rate of liquid HAN is slow. HAN is unsuitable for monopropellant catalytic decomposition engines because slowly decomposing HAN cannot decompose fast enough for sufficient engine efficiency and rapidly decomposing HAN results in excessively high temperatures, for example, greater than 2000° F., that are known to damage catalyst materials. As such, HAN is not suitable for decomposition propulsion systems. Two difficulties in designing a HAN based propulsion system are achieving the desired ignition characteristics and maintaining acceptable hardware temperatures during operation. These and other disadvantages are solved or reduced using the invention.